Gas turbine engines, such as those used to power modern aircraft or in industrial applications, include a compressor for pressurizing a supply of air, a combustor for burning a hydrocarbon fuel in the presence of the pressurized air, and a turbine for extracting energy from the resultant combustion gases. Generally, the compressor, combustor and turbine are disposed about a central engine axis with the compressor disposed axially upstream of the combustor and the turbine disposed axially downstream of the combustor. Air drawn into the engine passes axially through the compressor into the combustor wherein fuel is combusted in the air to generate and accelerate combustion gases that pass through the turbine and out the exhaust nozzle of the gas turbine engine. The combustion gases turn the turbine, which turns a shaft in common with the compressor to drive the compressor.
As the hot combustion gases pass through the turbine, various turbine elements, such as the turbine stator vanes and turbine rotor blades of the turbine, are exposed to the hot combustion gases, which may also be corrosive to the material of which those turbine elements are made. In order to protect the turbine elements from oxidation and corrosion due to exposure to the hot combustion gases, it is conventional practice to coat various turbine elements with one or more layers of a protective coating or coatings. For example, it is known to coat turbine stator vanes in gas turbine engines with an aluminide during the process of manufacturing the turbine modules.
The turbine of the gas turbine engine is generally an axially extending assembly of a plurality of turbine modules mounted to a shaft. Each turbine module may include one or more turbine stages. Each turbine stage includes a row of stationary airfoils, referred to as the stator vanes, and a row of airfoils, referred to as rotor blades, mounted on a rotor disk driven by the airflow passing over the rotor blades. The turbine may include a high pressure turbine including a plurality of high pressure stages in one or more modules assembled to a common shaft with a high pressure compressor, as well as a low pressure turbine including a plurality of low pressure stages in one or more modules assembled to a common shaft with a low pressure compressor and/or fan.
In the handling of the turbine modules during shipping, assembling of the turbine, and disassembling of the turbine for servicing or overhaul, the protective coating on the turbine elements may be damaged in local areas for example nicked, scrapped, cracked, scored or otherwise removed thereby exposing the base metal of which the turbine element is composed. If the damage is deemed sufficient to warrant repair, it is customary to repair the damaged coating by removing the module including the damaged element and disassembling the module so that the assembly having the damaged element may be replaced or repaired.
To repair a damaged area of coating on a turbine stator vane according to conventional practice, it is common to disassemble the turbine module to remove the stator assembly containing the damaged vane for service. The coating may then be stripped from the damaged vane, at least in the region surrounding and including the area of damaged coating, the remainder of the damaged vane exclusive of the stripped area is then masked, and a new coating is then applied to the area of damaged coating and the stripped area. The stator assembly is then placed in a furnace or oven at a desired temperature for a desired period of time to cure and heat treat the newly applied coating. The turbine module is then reassembled with the repaired stator assembly. Although effective for repairing the damaged turbine element, the process is time consuming and labor intensive as the turbine module must be dissembled to affect the coating repair since the turbine module itself is too large to be placed in a conventional heat treatment furnace or oven. Further, even if a furnace or oven were large enough to accommodate an entire turbine module, the whole surface of the turbine module exclusive of the stripped area to which the new coating has been applied would need to be masked to reduce the risk of contamination of the undamaged surface during the heat treatment process.
U.S. Pat. No. 6,560,870 discloses a method of applying a diffusion metal coating to a selective area of a turbine engine component having a deficiency of metal coating. To apply the diffusion metal coating in accord with the disclosed method, a metal source containing tape is positioned in contact with the selective area and held in contact with the selective area using a tape holder that is stable at high temperatures while the selective area is heated to an effective temperature and an effective amount of time to form a metal coating of predetermined thickness on the selective area. In the disclosed embodiment, a quartz infrared lamp is used to heat the selective area to a coating temperature of about 1800 F to about 2000 F under an inert atmosphere for about 3 to 8 hours.
U.S. Pat. No. 7,115,832 discloses a portable, hand-controlled microplasma spray coating apparatus that can be transported to on-site locations in the field to apply ceramic and metallic coatings to a variety of workpieces, including gas turbine engine parts. However, the use of such a microplasma spray coating apparatus to directly spray a plasma coating onto a base material is not generally satisfactory for application of coatings that require some degree of diffusion of the coating material into the base material to be effective.